Command system of missile guidance



1964 E. NORTON ETAL 3,156,435

COMMAND SYSTEM OF MISSILE GUIDANCE Filed Aug. 12, 1954 5 Sheets-Sheet lr T PREDICTED ,8

POSITION A7 I INTERCEPT %M/ss/L I6 35 26 I o o 24 TRANS.

REC. f/ ,M/SS/LE TARGET f MODUL- TRACK TRACK 2 NOR RADAR RADAR 20-/COMPUTER //4 I auRs'r ORDER (no) /56 TIME OF FLIGHT mw OROERET-PRED/CTED 7o ROs/r/O/v 5y mm: 4 DIV/DER mw ERROR M ,64 GVRO AX/$COORDINATE *MTMN 72 H. RESOLVER w E E 5 DIV/DER PITCH H X r ERROR PITCHORDER COMPARATOR [ST/CS 40 XMiIHM XTVTHT 30 22 f 1 PRED/CTER PREDICTERMISS/LE TARGET POSITION RESOLVER POSITION (3 cOOR.) (a com.)

ATTORNEY Nov. 10, 1964 E. NORTON ETAL 3,156,435

COMMAND SYSTEM OF MISSILE GUIDANCE Filed Aug. 12, 1954 3 Sheets-Sheet 2FIG. 2

' E. L. NORTON JW. SCHAEFER Nov. 10, 1964 NORTON ETAL 3,156,435

COMMAND SYSTEM OF MISSILE summer;

5 Sheets-Sheet 3 Filed Aug. 12, 1954 wt? Sin 6 V R E T 0 MM w A m LS m Sv, R m B N w m United States Patent COMMAND SYSTEM ill? MlSSlLE GUIDANCEEdward L. Norton, @ummit, and .lacoh W. Schaezfer,

Plainiieid, NJL, assignors to Bell Telephone Laboratories, incorporated,New York, NFL, a corporation of New York Filed Aug. 12, 1954, Ser. No.449,396 12 Qlaims. (Ql. 244l4) This invention relates to guided missilesystems and more particularly to systems for the effective control ofsupersonic antiaircraft missiles.

With the advent of high altitude, high speed bombing aircraft, it hasbecome increasingly diflicult to provide an adequate defense withconventional antiaircraft artillery. The extreme ranges at whichengagement must oc our to prevent a successful bombing attack and thehigh speeds attained by the aircraft involved result in undesirably longtimes of flight for even the highest velocity projectiles. Theprediction of future position of the target as required by the firecontrol problem is possible only when it may be assumed that the targetcourse will not change significantly during the flight of theprojectile. Obviously this assumption is not valid when long times offlight are involved. As a result it has been proposed to employ missileswhich can be steered while in flight as a defense against aircrafttargets. The time of flight of the missile is not a critical factor ifthe missile may be controlled after launching by means responsive to themaneuvers of the aircraft target to insure in erception of the targetdespite evasive action. Such systems necessarily include some means ofdetermining the present position of the target with respect to that ofthe missile and additional means for computing from this informationquantities which may be applied to the steering means and/or to thepower plant of the missile to correct its flight so that interception ofthe target may be accon1- plished.

Various systems of missile guidance have been proposed. One of the mostcommon of these is that in which the missile is equipped with a homingdevice. In such systems the missile carries some form of detectionapparatus which acts to locate the target and to provide informationfrom which a computing system also carried in the missile may generatesuitable control quantities to correct the missile path for interceptionof the target. Another system involves the so-called beam riding missilewherein a tracking device such as a Searchlight or a ground based radaris employed to track the target and the missile is launched in such aWay as to intercept the beam. Detection devices aboard the missilerespond to the beam and provide inputs for a computer also carriedaboard the missile which in turn generates suitable orders thereafter tomaintain the missile in the center of the beam.

These and similar systems of missile guidance all suffer from the basicdisadvantage that they cause the missile to follow an inefficienttrajectory, materially limiting its effectiveness. Further, the complexcontrol system required for guidance must be carried aboard eachmissile. Since the missiles are necessarily expendable, the cost of suchsystems becomes almost prohibitive. in addition the presence ofextensive control equipment aboard the missile detracts materially fromthe pay load (usually a warhead) and from the performance andreliability of the missile.

It is an object of the present invention to overcome these disadvantagesby minimizingthe amount of control ice in accordance with commandsissued by a ground based guidance equipment which is capable ofoperating to a high degree of precision and may be used repeatedly forthe control of any number of successive missiles. Both the target andthe missile are individually tracked by precision radars and data as tothe present position of both are applied to a computer which predictsthe future position of the target at a predetermined time (time ofintercept) and generates orders for transmission to the missile tocontrol the course thereof in such a way as to insure interception ofthe target at the predetermined time. Equipment aboard the missileprovides a frame of reference traveling with the missile andidentifiable at the location of the ground equipment so that properresponse to the commands may be obtained.

The above and other features of the invention will be described indetail in the following specification taken in connection with thedrawings, in which FIG. 1 is a block schematic diagram of the completecommand missile guidance system of the invention;

FIG. 2 is a diagram in schematic form of the missile employed in thesystem of FIG. 1 showing the equipment required aboard the missile; and

FIG. 3 is a vector diagram illustrating the missile command problem andthe manner in which the orders for transmission to the missile arecomputed.

In the broadest sense the antiaircraft guided missile system of theinvention comprises a target tracking device It), a missile trackingdevice 12, a computer id and a missile 16, all as shown in the blockdiagram of FIG. 1. Target tracking device ill which preferably comprisesa precision automatic tracking radar is arranged continuously to providedata as to the present position of a target aircraft 1%. This trackingradar may, for example, be similar to the well known SCR-584 radar whichis described in detail in Electronics for November 1945 beginning atpage 104, for December 1945 beginning at page 104 and for February 1946beginning at page 110. Briefly this radar is an automatic tracking radaremploying conical lobing whereby the radar antenna may be causedcontinuously to track the target in both elevation and azimuth and toproduce electrical representations of these quantities measured withrespect to predetermined reference quantities. In addition this radarincludes a range unit also responsive to the reflected radar pulseswhich automatically maintains itself adjusted to represent the slantrange to the target.

The output of target tracking radar 10 then comprises three quantities,namely, the slant range from the location 0 of the tracking radar to thetarget, the azimuth of the target measured from a predetermineddirection and the elevation of the target measured from a predeterminedreference (normally the horizontal plane). The data units associatedwith this particular radar vary with the application to be made of theavailable information. It is assumed for the present purposes that themechanical shaft positions corresponding to azimuth, range and elevationare converted to electrical quantities for individual transmission overa connector 2% to a predictor 22. The form in which the target positiondata are determined is not significant since the means for convertingdata in one coordinate system to another system are well known in thecomputer art.

The missile tracking device 12 may be similar to target tracking devicelid and may in the same way produce output quantities proportional tothe slant range, the elevation, and the azimuth of the missile asmeasured at the location of the missile tracking device. As shown inFIG. 1, however, certain advantageous modifications have been made inthe missile tracking device to improve the performance thereof.Basically these modifications involve recognition of the fact that theantiaircraft missile w presentsan extremely difficult target for atracking radar. Accordingly it has been found desirable to employ asocalled radar beacon system rather than a conventional radar. For thispurpose pulses are radiated from a transmitter 24 at a radio frequencyh. The missile, as will be explained in greater detail hereinafter,carries a responser which is responsive to pulses of frequency f andradiates pulses of frequency f These pulses are picked up by antenna 25and directed to a receiver 26 responsive to that radio frequency. Thereceiver 26 operates in a manner identical to that of target trackingradar It to provide the required output quantities for transmission overa connector 28 to a second predictor 30.

Radar beacon systems of the type contemplated for use in the missiletracking system are Well known in the art and are discussed in detail inRadar Beacons by Roberts, Vol. 3 of The Radiation Laboratory Series,McGraw-Hill, 1947. Modification of the SCR-S 84 radar referred to aboveas illustrative of the ground based missile tracking device for thistype of performance may easily be accomplished merely by tuning thereceiver to frequency f rather than f The missile-borne equipment willbe considered hereinafter.

As assumed above, the quantities applied to predictor 22 indicate thepresent position of target 18 in spherical coordinates with respect tothe location of the target tracking radar While those applied topredictor 3t) represent similar information as to the position of themissile with reference to the location of the missile tracking radar.For ease in computation it is considered desirable to convert thisinformation into rectangular coordinates With the origin at the locationof the target tracking radar. Such coordinate conversion is Well knownin the art and may be accomplished as described, for example, in Patent2,408,081 to 'Lovell et 211., September 24, 1946. Conveniently eachpredictor 22 and 30 includes a coordinate converter acting to convertinput quantities to rectangular coordinates (X, Y and H, where X and Yare orthogonal axes in the horizontal ground plane and H is the verticaldistance from the XY plane) with origins at the locations of therespective tracking radars. The necessary offset or parallax correctionsrequired to convert the data as to missile position to the coordinatesystem having its origin at the location of the target tracking radarmay be set in manually as potentials of suitable polarity along thethree rectangular coordinates. These corrections are constants and oncedetermined at the time at which the two tracking radars are emplaced,need not be changed unless the emplacement of the guidance equipment ischanged.

Computer 14 which includes predictors 22 and 36 acts on the basis of anassumed time of flight for the missile to reach a point of interceptionwith the target. Using this time of flight, the future positions of thetarget and missile at the predicted time of interception are computedindependently. These positions are compared coordinate by coordinate tofind the predicted position error at the predicted time of interception.The corrections ineither or both the time of flight assumed at the'outset and the course of the missile required to insure 4.- untilinterception). In each predictor the present position data aredifferentiated to obtain velocities and these velocities are multipliedby the time of flight to obtain future position. These operations arecarried out in each instance in terms of components along the orthogonalX, Y and H coordinates which are ground coordinates with the origin atthe location of the target tracking radar as indicated in FIG. 3 of thedrawing.

As in the reference patent each predictor 22 and provides threecoordinate output data representing the predicted position of the targetor the missile as the case may be when the predicted time of interceptoccurs. These quantities are identified in FIG. 1 as X Y and H for thetarget and X Y and H for the missile.

These quantities are applied to a comparator 32 in which they aresubtracted coordinate by coordinate to obtain predicted position errorquantities. The error outputs shown at the output of comparator 32 are EShOWll'lg the position error in the H direction, and E and E, showingthe position errors in the X and Y directions, respectively.

A better understanding of the significance of these quantities may beobtained by reference to the vector diagram of FIG. 3. Here the presentheading of the missile is represented by the arrow labeled h and that ofthe target by the arrow 12,. It will be assumed that at the predictedtime of intercept (T :0) the missile will have reached a position aspecified with respect to the X, Y and H axes by the quantities X Y andH referred to above and the target a position b similarly speci fled bythe quantities X Y and H The total position error at T=0 for theseassumptions is measured by the vector E, the components of which E E andE in the X, Y, and H direction respectively are as shown in FIG. 3. Itis apparent that these quantities are measured with reference to a setof coordinates fixed with respect to the target tracking radar and donot indicate directly what maneuvers must be performed by the missile toassure interception. It is necessary, therefore, to convert these errorquantities into quantities measured with respect to a frame of referencetraveling with the missile The method of providing the required frame ofreference traveling with the missile and still identifiable at thelocation of the guidance equipment will become apparent with referenceto the diagram of FIG. 2. Here the missile 16 is shown as carrying aso-called free-free gyroscopev 34, the rotor of which is suspended in aconventional gimbal system. The outer gimbal 36 is journaled forrotation about the longitudinal axis of the missile While the innergimbal 38 is journaled in the outer gimbal for rotation about an axisnormal to the longitudinal axis of the missile. The gyroscope rotor isin turn journaled in the inner gimbal and spins about an axis normal tothe inner gimbal axis. This third axis is hereinafter referred V to asthe gyro spin axis. In accordance with Well underinterception may bedetermined from this information.

fired. I

Assuming a time of flight, predictors 22 and latl determine thefurturepositions of both missile and target at r the predicted. time ofintercept (T=.0 Where T is the time case some measure of control remainsafter the missile is stood principles the gyroscope acts to maintain thegyro spin axis G at a fixed orientation in space regardless of themaneuvers of the missile in which the gyroscope is mounted. I

Conveniently, the gyro spin axis is initially oriented in the XY(horizontal) plane and, preferably but not necessarily, normal to theplane of the initial trajectory to the target. This axis and thelongitudinal axis of the missile define a reference plane attached tothe missile and'identiliable at all times at the location of theguidance equipment as will be pointed out below. 7 V

The initial orientation of the spin axis may be determined before themissile is launched and the heading (orientation of the longitudinalaxis) 'ofthe missile at any time may be obtainedifrom the missilevelocity vector .Which .is determined as one. of the necessary functionsof predictor iitl. ltwillfbefrecalled that the rates of change of eachof the X, Y and H quantities representing the present position of themissile are determined in the prediction process. These components ofthe missile velocity may be combined to give a quantity proportional tothe missile velocity in the direction of the missile flight path. Thiscombination of velocity components is performed by missile headingresolver 49 which is a conventional coordinate resolution device of thetype employed generally in fire control computers. It is noted that theframe of reference traveling with the missile is based upon thegyroscope spin axis and the longitudinal axis of the missile while theinformation available to the computer includes the orientation of thegyro spin axis and the missile velocity. It has been found that, becauseof the continuous control of the missile afforded by the command systemof the invention, the missile velocity vector may be considered to havethe same orientation as the longitudinal axis of the missile. Any errorso introduced is within the range of correction of the command system.It will be understood, however, that if sufiicient information isavailable as to the aerodynamic performance of the missile, additionalcomputation equipment may be provided to determine the orientation ofthe longitudinal axis of the missile from the available missile velocityinformation.

It now becomes necessary to steer the missile with respect to the frameof reference just considered. As shown in FIG. 2 the missile is providedwith two sets of paired steering fins 4t) and 42, respectively. Pins 40are mounted on a shaft 44 normal to the longitudinal axis of the missileand fins 42 are mounted on a shaft 46 which is normal to both thelongitudinal axis of the missile and the shaft 44. Shafts 44 and 46 thusconstitute a pair of steering axes which will be referred to as the yawand pitch axes, respectively, and which control the orientation of themissile in two orthogonal planes. As a matter of convenience the planenormal to shaft 44 will be referred to as the yaw plane (which is thesame as the reference plane considered above) and that normal to shaft46 as the pitch plane (which includes the longitudinal axis of themissile and is normal to the reference plane). Fins then constitute theyaw steering fins and fins 42 the pitch steering fins. These fins arepositioned in response to steering orders transmitted from computer 14to control the course of the missile after launching.

It is apparent that the course of the missile can be properly controlledonly if the steering planes defined above remain properly oriented withrespect to the plane of reference considered above and defined by thegyro spin axis and the longitudinal axis of the missile. This conditionrequires roll stabilization of the missile. For this purpose the missileis equipped with paired cruciform tail fins 48 and 5t? and adjustableailerons 52 are provided upon the trailing edges of at least one pair oftail fins (48 in FIG. 2). Conveniently gyroscope 34 is employed tocontrol ailerons 52 in such a way as to roll stabilize the missile withrespect to the reference plane.

The outer gimbal 36 of the gyroscope 34 is coupled to a pick-off orother sensing device 54 through a shaft which forms an extension of theouter gimbal axis. The pickoff or sensing device 54 comprises a sourceof error signal for a servo system including an amplifier 56, a motor 58and suitable linkage 6t 62 whereby opposite deflections of upper andlower ailerons 52 may be produced by motor 5 3. Although these elements,together with the gyroscope 34 which constitutes the error detectingelement, may be connected to form any of a large number of known kindsof servo system, it may be assumed for the purposes of the presentdescription that a simple direct current servo system is employed.

For such a servo system pick-off device 54 comprises a potentiometerprovided with a center-tapped winding to which direct current potentialsare applied from a source such as a battery (not shown to avoid unduecomplexity in the drawing). The circuit is so arranged that no output isproduced when shaft 44 is normal to the reference plane. Whenever thepotentiometer arm is moved from this normal (null) position an output isdeveloped, the amplitude and polarity of which are indicative of theamount and direction of the roll of the missile from the desiredposition. After amplification this output may control motor 58 as in theusual direct-current servo sys tem, causing deflection of the aileronsin the proper direction to return the Inissile to the desired.orientation as indicated by the null output from potentiometer 54.

It will be understood that as a result of such roll stabilization fins40 are effective to produce steering forces in the yaw plane and theother pair of steering fins 42 act to produce steering forces in thepitch plane. Thus there is provided a set of orthogonal referencecoordinates traveling with the missile and comprising the yaw axis y,the pitch axis 1 and the missile heading h Further the orientation ofthis reference system in space is continuously determinable at thelocation of the ground guidance equipment.

It will be recognized that by the usual process of coordinate conversionthe total position error E shown in FIG. 3 may be expressed in terms ofcomponents along the reference axes traveling with the missile. Suchconversion may be accomplished in coordinate resolver 64, FIG. 1 asoutlined beginning page 279 of Electronic Analog Computers by Korn andKorn, or in Patent 2,658,674 to Darlington et al., November 10, 1953,particularly in FIG. 38 and the specification beginning at column 92thereof. This coordinate resolver accepts the three position errorcomponents E E and B shown in FIG. 3 and in addition accepts quantitiesfrom the output of missile heading resolver 40 indicating theorientation of missile heading axis in space and quantities representingthe positon of the gyro spin axis G which may be set into the coordinateresolver as constants. The outputs of coordinate resolver 64 are E, andli representing position errors in the yaw and pitch planes and measuredalong the pitch and yaw axes respectively and 18,, representing aposition error along the missile path. These components are shown inFIG. 3 with respect to the missile axes p, y and h (drawn with an originat point 0, the present position of the missile).

E, the predicted position error along the missile path is a measure ofthe adjustment which must be made in the time of flight (or the speed)ofthe missile to cause interception of the target. If it be assumed thatthe speed of the missile is not subject to external control once themissile is launched, this correction must be made by varying the time offlight originally assumed at the outset of the computation process andemployed as one input to each of predictors 22 and 30. This quantity is,therefore, applied to a time of flight servo mechanism 66 and controlsthe setting of input quantities applied to the two predictors possiblyas a shaft rotation as in the predictors shown in Patent 2,408,081,referred to above. Since the speed of the target is assumed to beconstant, the quantity controlling predictor 22 is applied directlythereto. However, the corresponding quantity for predictor 36 which isassociated with the missile section of the computer is modified inaccordance with the ballistic characteristics of the missile beforeapplication to predictor 30. Such modifications are accomplished byapparatus as wherein appropriate changes are made in the value of timeof flight. These changes are ordinarily accomplished by adding bothfixed and variable components to that corresponding to the time offlight as shown for example in' FIG. 8A of Patent 2,408,081 to whichrefererence has been made above.

It will be understood from the above that the necessary functions forthe continous prediction process performed by the computer are providedby way of the time of flight servo mechanism and the target and missiletracking radars. The pitch and yaw error outputs E and E from mounted.-

coordinate resolver 64 are employed for the generation of steeringorders for the missile. These quantities depend of course upon the timeof flight fed back to predictors 22 and 30 as discussed above. As amatter of convenience in control of the missile it has been founddesirable to convert these position orders into acceleration orders, is.the quantities representing the position errors measured along the yawand pitch axes are converted intov lateral accelerations in the pitchand yaw planes, respectively, such that the missile will be at the pointof predicted interception at the predicted time of interception. For thegeneration of such orders the quantities E and E are appliedrespectively to dividers 70 and 72 in which each -isdivided twice by thetime of flight produced as the output-'of nnit'd 'The yaw and pitchorders appearing at the outputs of dividers 70 and 72 respectively arethus proportional to the accelerations requiredto cause the missile toreach the predicted point of interception at the predicted time ofinterception. These orders may be transmitted to the missile by anyconvenient means, for example, by a high frequency radio communicationchannel.

Alternatively and as shown in FIG. 1, the acceleration orders for themissile are transmitted by modulation of the repetition rate of themissile tracking pulse transmitter. The necessary control quantities maybe transmitted on a time division basis or any other convenient basis bythe action of a modulator 74 associated With transmitter 24. Varioussystems of signaling over the radar beam are described in Section 11.2of Radar Beacons, vol. 3 of the Radiation Laboratories Series. Accordingto one such system the two control signals are transmitted as audiofrequency signals frequency modulated upon the pulses from the tracktransmitter. Either one or both of the frequencies can thus betransmitted depending upon the steering orders required at a particulartime, the modulating wave comprising either one or the sum of the audiofrequencies.

Also transmitted to the missile and conveniently by interruption of-allmodulation upon the radar beam is the so-called burst order which at atime related to the predicted' time of intercept causes the warhead ofthe missile to explode. Ordinarily this order is transmitted a fewmicroseconds prior tothe time when T =0.

The remaining equipment carried aboard the missile may now be.considered. As has been stated above the missile carries a transponderresponsive to pulses from the ground based missile tracking equipment.This transponder includes a receiver 76' tuned to frequency h associatedwith antenna 78, and a microwave pulse transmitter 80 which is triggeredby the output of receiver 76 and which radiates pulses of frequency ffrom antenna 82. These pulses when received at the location of theguidance. equipment permit tracking of the missile.

Radio receiver 76 also serves to receive the various orders transmittedfrom the guidance equipment and intended to control the steering fins ofthe'missile and the burstingv of the warhead-at appropriate times.Accordingly it is provided with demodulation and channel separatingequipment appropriate to the nature of the modulation and multiplexsystems employed to transmit these orders. In any event the receiver isdesigned with reference to the particular transmitter 24, FIG. 1,employed in the guidance equipment and produces three output signalscorresponding respectively to the pitch and yaw acceleration orders andthe warhead burst order.

As shown in FIG. 2', each pair'of missile steering fins is driven by aservo motorgin response to they appropriate orders occurring atthe'output of receiver 76. Thusa motor 84 is geared to the shaft 44 uqn' which steering fins 40 are mounted and a motor 86 is geared to thecorresponding shaft' 46 upon which steering fins 42 are It will berecalled that the missileis to be made responsive to accelerationorders, Accordingly for each set of'steering this the appropriate orderappearing at the output of radio receiver 76 is applied to an amplifierto which is also applied an output of an accelerometer. These twoquantities are applied in opposition and the motor is driven from theoutput of the amplifier until the accelerometer indicates that thedesired lateral acceleration has been introduced. When such a conditionoccurs the output of the amplifier is reduced to zero and the motorstops. If the lateral acceleration increases, the output of theaccelerometer exceeds the order output of the receiver and the motor isdriven in the appropriate direction to return the acceleration to therequired value. The yaw steering fins 40 are thus controlled by theoutput of the comparison amplifier 88 to which is applied the output ofan accelerometer 9t) oriented in the missile to indicate accelerationsin the yaw plane. Simiiarly' the pitch steering fins 42 are controlledby a comparison amplifier 92 to which is applied the pitch order outputof the receiver and the output of an accelerometer 94 oriented in themissile to detect accelerations in the pitch plane.

The remaining equipment in the missile includes a warhead 96 furnishedwith an appropriate detonator which may be actuated by an electricimpulse known as the burst order received from the appropriate output ofreceiver 76 and applied to the warhead over lead 98;

In summary, the command system of missile. guidance according to theinvention minimizes the amount and the cost of the equipment whichtravels with the missile and is expended with each missile and placesthe majority-of the precision control equipment on the ground where itcan be used repeatedly and where adequately controlled environment formaximum efiiciency of operation can be provided. In the operation of thesystem, a target is tracked by the target tracking radar and when thecomputer circuits have had time to settle and the predicted point ofinterception is Within range of capability of the missile, a'missile islaunched. As the missile is launched it is tracked by the missiletracking equipment and the computer continuously determines correctionsto be made in the missile course to insure interception of the target ata predicted time of interception. So long as the lateral accelerationsof which the missile is capable in the course of maneuvers exceed thoseof the'target aircraft, interception of the target can be accomplishedregardless of evasive maneuvers attempted thereby. When the predictedtime of interception approaches, the computer produces. a burst orderwhich actuates the Warhead of V the missile at such a time as toinsurethe maximum effectiveness against the target of the resultingdetonation. v

What is claimed is:

1. In an antiaircraft system a missile, means in said missile forestablishing a reference axis of fixed orienta tion with respect-to theearth, means for steering said missile about a pair of orthogonal axesfixed with respect to the missile, means for stabilizing said missile tomaintime and means for transmitting control signals based on' f saidpredicted paths to the missile steering means to insure interception ofthe target by the missile at said time.

2. In an antiaircraft system a missile, means in said missile forproducing lateral accelerations thereof in a pair of orthogonalacceleration planes fixed with reg spect to said missileandintersectingin a line coincident with the'longitudinal axis ofsaidmissilepmeans aboard.

said missile for establishing a reference axis of fixed orientation withrespect to vthe earth, means for maintaining said acceleration planes infixed relationship to the reference plane defined by said referenceaxisandthe longitudinal axis of said missile, means for continuouslyestablishing the present positions of said missile and a target aircraftin space, means for computing from said present positions and an assumedfuture time of interception corrections in the course of said missile toinsure interception of the target at said time and means fortransmitting orders from said computer to said acceleration producingmeans in said missile correspondingly to produce lateral accelerationsin said acceleration planes to correct the missile course forinterception of the target at said time.

3. In an antiaircraft system a self-propelled missile, a gyroscopemounted in a gimbal system in said missile With the outer gimbal axiscoincident with the longitudinal axis of said missile, means forsteering said missile about a pair of axes normal to the longitudinalaxis of the missile and initially having a predetermined positionalrelationship to the reference plane defined by said longitudinal axisand the spin axis of said gyroscope, means responsive to rotation of theouter gimbal with respect to said steering axes to rotate the missileabout its longitudinal axis for the purpose of maintaining said initialrelationship, means for continuously establishing the positions of saidmissile and of a target aircraft in space, means for computing from saidpositions future positions of said missile and said target aircraft anda course for said missile to insure interception of the target aircraftat a future time and means for transmitting control signals dependentupon the predicted missile course to said steering means.

4. in an antiaircraft system a missile, means in said missile forestablishing a reference axis of fixed orientation with respect to theearth, means for steering said missile about a pair of axes fixed withrespect to said missile and normal to the longitudinal axis thereof,ailerons for controlling the roll of the missile about said longitudinalaxis, means for sensing rotation of the missile about said longitudinalaxis with respect to said reference axis, means responsive to saidsensing means for deflecting said ailerons to return the steering axesto a predetermined relationship with respect to a reference planedefined by said longitudinal and reference axes, means at a groundcontrol station for determining from the present positions of saidmissile and of a target aircraft, means for determining from saidpresent positions a missile course to insure interception of said targetaircraft, and means for transmitting orders from said ground station tosaid missile to produce maneuvers of missile about said steering axes tomaintain the course to interception of said target.

5. In an antiaircraft system a self-propelled missile, means fortracking a target and producing quantities representative of itsposition in rectangular earth coordinates having their origin at thelocation of said tracking means, missile tracking means producing outputquantities representative of the position of said missile in said earthcoordinates, means in said missile for establishing a refer ence axis offixed orientation with respect to said earth coordinates, steering meansaboard said missile for producing lateral accelerations thereof inorthogonal planes bearing a fixed relationship to said missile, meansfor roll stabilizing said missile to maintain said acceleration planesin fixed relationship to said reference plane, means for computing fromthe positions of said missile and said target in said earth coordinatescorrections in the course of said missile required to insureinterception of said target by the missile at a future time, means forconverting said course corrections into acceleration orders related tosaid reference plane aboard the missile and means for transmitting saidorders to said missile to control lateral accelerations thereof in saidacceleration planes.

6. in an antiaircraft system a self-propelled missile, means in saidmissile for establishing a reference axis of fixed orientation withrespect to the earth, means for producing lateral accelerations of themissile in a pair of acceleration planes intersecting along thelongitudinal axis of the missile, means for stabilizing the missile tomaintain a fixed relationship between said acceleration planes and thereference plane defined by said reference axis and the longitudinal axisof the missile tracking means for continuously establishing the positionof a target in space, other tracking means for similarly establishingthe position of said missile, means for computing from the target andmissile position information position errors at a predicted time ofinterception as measured in said reference plane and said accelerationplanes aboard the missile, means for converting said errors intoaccelerations of the missile required to eliminate the errors by thepredicted time of interception, and means for transmitting ordersrepresenting said accelerations to said lateral acceleration producingmeans aboard the missile to correct the course thereof.

7. In an antiaircraft system a missile, means in said missile forestablishing a reference axis of fixed orientation with respect to theearth, means for steering said missile about a pair of axes fixed Withrespect to the missile, means for stabilizing the missile to maintainsaid steering axes in fixed relationship at all times to the planedefined by said reference axis and the longitudinal axis of the missile,means for determining the present positions of said missile and of atarget, means for predicting from the present position of said targetthe position thereof at a chosen time of interception, means forpredicting from the present position of said missile and the ballisticcharacteristics of the missile the position of the missile at said time,means for determining from said predicted positions a position error incomponents along a set of axes including the longitudinal axis of saidmissile and said steering axes, and means for producing from at leastsome of said position error components control signals for transmissionto said missile steering means to eliminate said position errors by saidtime.

8. In an antiaircraft system a self-propelled missile, means in saidmissile for establishing a reference axis of fixed orientation withrespect to the earth, means for producing lateral accelerations of themissile in a pair of acceleration planes intersecting along thelongitudinal axis of said missile, means for stabilizing said missile tomaintain said acceleration planes in fixed relationship to the referenceplane defined by said reference axis and the longitudinal axis of saidmissile, ground based means for determining from quantities continuouslyrepresentative of the present positions of said missile and of a targett e relative positions thereof at a predetermined time of interception,means responsive to the relative positions at said predetermined time toobtain position errors in terms of earth coordinates having an origin atthe location of said ground based means, means for converting saidposition errors in earth coordinates to corresponding errors in terms ofa reference system based upon said reference and said accelerationplanes and traveling with said missile, means for determining from saidcorresponding errors the requisite accelerations in said accelerationplanes to eliminate said errors at said predetermined time ofinterception, and means for transmitting signals proportional to saidaccelerations to said missile to control said missile steering means toproduce such accelerations.

9. In an antiaircraft system a self-propelled missile, a Warhead in saidmissile and detonating means therefor, means also in said missile forestablishing a reference axis of fixed orientation with respect to theearth, means for steering said missile about a pair of axes fixed Withrespect to said missile, means for stabilizing said missile to maintainone of said steering axes normal at all times to the plane defined bysaid reference axis and the longitudinal axis of said missile, means forcontinuously establishing the positions of said missile and of a targetin space, means for computing from said missile and target positioninformation a predicted path for said missile to insure interception ofthe target at a future time, means for transmitting control signalsdependent upon the predicted missile path to said missile steering meansand means for transmitting a control signal to said detonating means atsaid future time of interception.

10. In an antiaircraft system a self-propelled missile having a warheadand detonating means therefor, means aboard said missile forestablishing a reference axis of fixed orientation with respect to theearth, means for steering said missile with respect to a set ofrectangular coordinates, including the longitudinal axis of the missileand an axis perpendicular to the plane defined by said reference axisand the longitudinal axis of said missile, tracking means forcontinuously establishing the present positions of said missile and of atarget aircraft, means for predicting from said present positions thepositions of said missile and said aircraft at a predetermined time ofinterception, means for producing position error quantities With respectto the reference coordinates traveling With.

said missile, means aboard the missile and responsive to saidtransmitted quantities effective to change the missile course to reducethe position errors substantially to zero at said time, and meansoperative at said time to transmit an actuating signal to said Warheaddetonating means.

11. In an antiaircraft system a missile, a gyroscope mounted in saidmissile in a gimbal suspension having inner and outer gimbal axes, meansfor supporting the outer gimbal axis in coincidence With thelongitudinal axis of said missile, a potentiometer fixed in said missileand driven by relative rotation of said outer gimbal and said missile tosense roll of the missile about its longitudinal axis, ailerons mountedon said missile to control rotations thereof about said longitudinalaxis, a servo system including said potentiometer as a sensing elementand driving said ailerons to maintain a fixed angular relationshipbetween said outer gimbal and the missile, a pair of steering axes onsaid missile normal to said longitudinal axis, steering fins on saidaxes and means responsive to orders from a ground based computeraccepting the present positions of said missile and of a target andinformation as to the orientation of the gyroscope spin axis to actuatesaid fins for guiding said missile to interception with said target.

12. In an antiaircraft system a missile, means in said 5 missile forproducing lateral accelerations thereof in a pair of orthogonal planesfixed with respect to said missile and intersecting in a line coincidentwith the longitudinal axis of said missile, means aboard said missilefor establishing a reference axis of fixed orientation with respect tothe earth, means for maintaining said planes in fixed relationship tothe reference plane defined by said reference axis and the longitudinalaxis of said missile, means for continuously establishing the presentpositions of said missile and a target aircraft in space, means forcomputing from said present positions and an assumed future time ofinterception the lateral accelerations of the missile in said orthogonalplanes to insure interception of the target at said time, means fortransmitting control quantities proportional to said accelerations tosaid missile, means aboard the missile for determining the lateralaccelerations in said orthogonal planes, means for comparing thetransmitted acceleration quantity for each of said planes with themeasured acceleration in the respective plane, and means for adjustingthe corresponding acceleration'producing means aboard said missile tobring to equality the transmitted acceleration quantity and the measuredacceleration for each of said planes.

References Cited in the file of this patent OTHER REFERENCES 7 Guidancefor Missiles, by White; Coast Artillery Journal; November-December,1946; pp. 18-22.

1. IN AN ANTIAIRCRAFT SYSTEM A MISSILE, MEANS IN SAID MISSILE FORESTABLISHING A REFERENCE AXIS OF FIXED ORIENTATION WITH RESPECT TO THEEARTH, MEANS FOR STEERING SAID MISSILE ABOUT A PAIR OF ORTHOGONAL AXESFIXED WITH RESPECT TO THE MISSILE, MEANS FOR STABILIZING SAID MISSILE TOMAINTAIN ONE OF SAID STEERING AXES NORMAL AT ALL TIMES TO THE PLANEDEFINED BY SAID REFERENCE AXIS AND THE LONGITUDINAL AXIS OF THE MISSILE,TRACKING MEANS FOR CONTINUOUSLY ESTABLISHING THE POSITION OF A TARGETAIRCRAFT IN SPACE, OTHER TRACKING MEANS FOR SIMILARLY ESTABLISHING THEPOSITION OF